Aircraft engine inlet having zone of deformation

ABSTRACT

An inlet assembly for an aircraft engine nacelle includes an inner barrel, an outer barrel radially spaced from the inner barrel, and a rear support for supporting the outer barrel relative to the inner barrel. The rear support includes at least one radially extending stiffener and at least one energy absorber defining a zone of deformation adjacent the inner barrel. The rear support is deformable in the zone of deformation in response to forces applied during a fan blade-out event to prevent fracture of the rear support.

RELATED APPLICATIONS

This application is a divisional of U.S. appln. Ser. No. 11/488,980filed Jul. 19, 2006, the content of which is incorporated by referencein its entirety.

FIELD OF THE INVENTION

The present invention generally relates to gas turbine engines, and morespecifically to an inlet assembly for an aircraft engine nacelle havinga rear support with an energy absorbing zone.

BACKGROUND OF THE INVENTION

One type of jet engine typically used on large commercial airliners is aturbofan jet engine. Such turbofan engines have a compressor, combustorand turbine, and include a fan mounted on the front of the engine. Thesefans, some as large as 10 feet in diameter, draw air into the engine.Some of the air is sent to the compressor and the combustor, while therest bypasses these components through ducts along the outside of theengine. The fans include fan blades that rotate at speeds up toapproximately 9,000 rpm during operation of the engine and are enclosedin a fan housing that radially surrounds the fan. Turbofan engines forairliners typically include a nacelle that at least partially surroundsthe engine and provides an aerodynamic shell to minimize drag. Typicalnacelles generally comprise an inlet assembly located in front of thefan, generally being attached to the fan housing and shaped to directair into the engine, a fan cowl that encloses the fan, and a thrustreverser adjacent the rear part of the engine.

The inlet assemblies for turbofan jet engine nacelles typically includean inner barrel, an outer barrel, a forward bulkhead, and an aftbulkhead which spans between the inner and outer barrel. Typically, theaft bulkhead is an annular metal plate that is designed to fractureduring the initial (albeit extremely unlikely) impact of a fan blade-outevent or during the subsequent loads from the windmilling effect of theunbalanced fan. During a fan blade-out event, the fan blade impacts thefan housing at a high rate of speed generating a large impact force thatis transmitted from the fan housing to the inlet assembly attached tothe housing. After a fan blade-out event, the engine is shut down butthe fan can continues to rotate or “windmill” as the plane is flown to adestination for repair. However, the fan is generally unbalanced andthus continued rotation of the fan can cause high loads on the fanhousing. The bulkhead thus generally is reinforced with a series ofstructural support members extending between the inner barrel and outerbarrel and connected to the bulkhead at spaced apart locations. Inexisting designs, the support members are spaced such that the inletremains intact with the bulkhead being held together by the supportmembers. The use of structural support members on the aft bulkheads ofprior art inlet assemblies, however, increases the weight and cost ofthe nacelle.

Accordingly, there is a need for a nacelle for a turbofan jet enginehaving an inlet assembly that addresses the foregoing and other relatedand unrelated problems in the art.

SUMMARY OF THE INVENTION

In general, one aspect of the present invention is generally directed toan inlet assembly for a turbofan engine nacelle comprising an innerbarrel, an outer barrel radially spaced from the inner barrel, and arear support for supporting the outer barrel relative to the innerbarrel. The rear support comprises at least one plate extending betweenthe inner barrel and the outer barrel for forming a closed axial end ofthe inlet assembly. The plate has at least one stiffener extendingradially across the plate and at least one energy absorber formed in theplate. The energy absorber defines a zone of deformation adjacent theinner barrel. The rear support is deformable in the zone of deformationin response to an applied force during a fan blade-out event to preventfracture of the rear support.

In another aspect, the present invention is generally directed to amethod of manufacturing an inlet assembly for an aircraft engine nacellehaving an outer barrel and an inner barrel. The method comprises formingat least one radial stiffener in a rear support and forming at least onezone of deformation in the rear support by forming an arcuatelyextending energy absorber. The energy absorber is located in the rearsupport such that the support is deformable in the zone of deformationin response to an applied force during a fan blade-out event to preventfracture of the rear support. The rear support is attached to the outerbarrel and the inner barrel to form a closed axial end of the inletassembly.

In yet another aspect, the present invention is generally directed to anacelle for a turbofan engine. The nacelle comprises an inlet assemblyhaving an inner barrel, an outer barrel radially spaced from the innerbarrel, and a rear support for supporting the outer barrel relative tothe inner barrel. The rear support comprises at least one annular plateextending between the inner barrel and the outer barrel for forming aclosed axial end of the inlet assembly. The plate has at least onestiffener extending radially across the plate and an energy absorberformed in the plate. The energy absorber defines a zone of deformationadjacent the inner barrel. The rear support is deformable in the zone ofdeformation in response to an applied force during a fan blade-out eventto prevent fracture of the rear support.

Various objects, features and advantages of the present invention willbecome apparent to those skilled in the art upon reading the followingdetailed description, when taken in conjunction with the accompanyingdrawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial perspective view of an aircraft having turbofan jetengines each with a nacelle according to a first embodiment of theinvention.

FIG. 2 is a perspective of an aircraft turbofan jet engine and nacellestructure.

FIG. 3 is a perspective of an inlet assembly of an aircraft turbofan jetengine nacelle, as viewed from a back-to-front perspective.

FIG. 4 is and end view of an enlarged portion of the inlet assembly FIG.3.

FIG. 5 is a cross-sectional view of the inlet assembly taken along theplane including line 5-5 of FIG. 4.

FIG. 6 is a partial cross-sectional view of an inlet assembly of asecond embodiment of the invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings in which like numerals indicate like partsthroughout the several views, the present invention generally relates toa nacelle, generally indicated at 1 in FIG. 1, for a turbofan jet engine3, mounted on an aircraft 5. As shown in FIG. 1, the nacelle 1 enclosesthe engine 3 and forms the outer aerodynamic covering of the engine. Asshown in FIG. 2, the turbofan jet engine 3 includes a fan assembly,generally indicated at 7, located towards the front of the engine andhaving a fan 9 having fan blades 11 rotatably mounted in a fan housing13 of the engine. It is understood that the nacelle 1 is only partiallyshown in FIG. 2 and includes an inlet assembly 17, mounted on the frontof the fan assembly 7 and a fan cowl 19, including doors 21, 23 thatenclose the fan assembly. It further is understood that the nacelle 1may include other components and systems not shown in the illustratedembodiments (e.g., thrust reverser and de-icing system) withoutdeparting from the scope of this invention.

As shown in FIG. 2, the inlet assembly 17 includes a leading edgesection 27 that includes a forward lip 29 shaped to provide a naturalairflow along the forwardmost surfaces of the engine 3. The inletassembly 17 has a rear body 31 that is connected to the rear of theleading edge section 27 and the front of the fan cowl 19, and includes arear support 35 (FIG. 3). The inlet assembly 17 further includes aninner barrel 39, forming an inner radial surface 40 of the inletassembly, and an outer barrel 41. The outer barrel 41 is radially spacedfrom the inner barrel 39 and forms an outer radial surface 42 of theinlet assembly. The outer barrel 41 is formed by adjacent outer radialwalls of the leading edge assembly 27 and the rear body 31. As shown inFIG. 5, the inner barrel 39 has a flange 43 at the rear of the rear body31 for connecting to a corresponding flange 45 on the fan housing 13.The connection of the inner barrel 39 to the fan housing 13 enablesimpact forces imparted against the fan housing to be transmitted to theinlet assembly 17 via the connected flanges 43, 45.

The rear support 35 forms a closed axial end of the inlet assembly 17and supports the outer barrel 41 relative to the inner barrel 39 asindicated in FIG. 3. As shown in FIGS. 3-5, the rear support 35generally is a plate 47 formed from or having multiple sections orsegments 18. Each section 18 of the rear support 35 has a series ofcircumferentially spaced ribs 61 extending radially in the plate toincrease the sonic resistance of the support 35 and stiffeners 63between the ribs 61 that increase the stiffness of the support 35.Energy absorbers 67 extend arcuately in the support 35 to form annularlyspaced zones of deformation 71 adjacent the inner barrel 39. In theillustrated embodiment, the ribs 61 each have an inner radial edge 73(FIG. 4) that is adjacent the radial outer edge 75 of an energy absorber67. In one embodiment, and as shown in FIG. 5, each rib 61 is formed bycreating a semicircular depression in the annular plate 47. Thestiffeners 63 comprise the flat surface area of the plate betweenadjacent ribs 61 and the energy absorber 67. In the illustratedembodiment, each segment 18 of the support 35 comprises four ribs 61,three stiffeners 63, and one energy absorber 67.

In the illustrated embodiment, and as shown in FIG. 5, each energyabsorber 67 is an arcuate rib comprising a semicircular depression inthe plate 47. At each zone of deformation 71, a respective energyabsorber 67 is deformable in response to an applied force transmittedfrom the fan housing 13 during a fan blade-out event to prevent fractureof the rear support 35. Each zone of deformation 71 acts as a crumplezone that absorbs the impact energy transmitted from the fan housing 13and allows the rear support 35 to deform in response to the energy.Because the rear support 35 deforms in the zone of deformation 71, therear support 35 is prevented from cracking or otherwise fracturing inresponse to the impact force from the fan housing 13 during a fan bladeout event and the resulting force imparted on the rear support 35 duringthe subsequent windmilling of the unbalanced fan 9. In the illustratedembodiment, each radial rib 61 and arcuate rib 67 is a semi-circulardepression, but it is understood that each rib may have othercross-sectional shapes such as curves having a varying radius ofcurvature and/or portions that are substantially flat.

In the illustrated embodiment, the rear support 35 has a materialthickness of approximately 0.080 inches and a radial length R1 (FIGS. 4and 5) of approximately 14 inches, and an inner radius R2 (FIG. 3) ofapproximately 57 inches. Each rib 61 has a radial length R3 (FIG. 4) ofapproximately 9 inches, a width W1 of approximately ¾ inch toapproximately 2½ inches, and a spacing W2 between adjacent ribs ofapproximately 6 inches. Each energy absorber 67 has a radial length R4(FIG. 4) of approximately ¾ inch to approximately 2½ inches with theouter radial edge of the energy absorber being spaced from the innerradial edge of the rear support 35 a radial length R5 of approximately 3inches. The energy absorbers 67 and ribs 61 have a radius R6 (FIG. 5) ofapproximately half the length R4 and may generally be in the range ofapproximately ⅜ inch to approximately 1¼ inches. The dimensionalinformation described herein is intended to illustrative of oneembodiment of the invention and should not be construed as limiting thescope of the invention because the dimensions of the invention may varyfrom the dimensions and ranges described herein without departing fromthe scope of this invention. In the illustrated embodiment, each zone ofdeformation 71 is defined by an angular section of the rear support 35including four ribs 61 and three stiffeners 63 connected by one energyabsorber 67 with each angular section having an angle A1 (FIGS. 3 and 4)of approximately 20 degrees. It is understood that the number ofstiffeners 63, ribs 61, and energy absorbers 67 of each zone ofdeformation 71 may vary and that the angle Al may be more or less than20 degrees without departing from the scope of this invention.

The multiple sections 18 of the rear support 35 are attached at an innerradial edge to the inner barrel 39 by fasteners 51 connecting the rearsupport 35 to a flange 53 mounted on the external surface of the innerbarrel 39 forward of the rear flange 43 connecting the inlet assembly tothe fan housing 13. The sections 18 of the rear support 35 are attachedat an outer radial edge to the outer barrel 41 by fasteners 55connecting the rear support 35 to a flange 57 mounted on inner surfaceof the outer barrel 41 at the rear axial end of the outer barrel 41. Theforces (e.g., impact forces, vibration forces, etc.) exerted on the fanhousing 13 during operation of the engine 3 are transferred to the innerbarrel 39 and to the rear support 35 attached to the inner barrel alongthe flange 53.

As shown in FIG. 4, adjacent sections 18 of the rear support 35 areconnected by fasteners 81 extending through overlapping adjacent edgemargins of the adjacent sections. Each section 18 of the rear support 35may be manufactured by conventional fabrication processes such as metalcasting or hydroforming with the curved radial ribs 61 and the curvedarcuate rib of the energy absorbers 67 preformed in the molding used tocast each section. In the embodiment of FIGS. 3-5, the rear support 35is made from 2024-T4 Aluminum but it is understood that the supportcould be made from other suitable materials and variations thereof(e.g., titanium, steel, CRES, Inconel, etc.). Material selection isbased on numerous factors including environmental conditions, weight,cost, and design conditions. As shown in FIG. 5, the rear support 35 isgenerally vertical between the inner barrel 39 and outer barrel 41. Itis understood that in alternative embodiments, the rear support 35 couldbe positioned at an angle up to approximately 45 degrees in the forwardor aft direction without departing from the scope of this invention.

The rear support 35 of the inlet assembly 17 of the present invention isdesigned to withstand the high stress and strain imparted to the inletassembly from the fan housing 13 during a fan blade-out event. During afan blade-out event, a fan blade 11 becomes partially or fully detachedfrom the fan 9 and may impact the fan housing 13 causing a large initialimpact force to be transmitted to the inlet assembly 17. As a result ofsuch impact, the fan 9 may become unbalanced and continue to rotate orwindmill during subsequent flight of the aircraft 5. The rotation of theunbalanced fan 9 exerts further stress and strain on the fan housing 13which is transmitted to the inner barrel 39 and rear support 35 of theinlet assembly 17. The zones of deformation 71 of the rear support 35are designed to absorb the forces from the fan housing 13 during a fanblade-out event without fracture of the rear support. The energyabsorbers 67 absorb the energy and allow the rear support 35 to deformin the zone of deformation 71 in response to the impact forces duringthe fan blade-out event. Deformation of the rear support 35 in the zoneof deformation 71 allows the rear support to withstand the fan blade-outevent without fracture of the support or failure of the fasteners 51, 55attaching the support to the inner barrel 39 and the outer barrel 41. Inthis way, the structural integrity of the rear support 35 is maintainedduring a fan blade-out event so that the inlet assembly 17 stays intactand no parts are broken off or discharged from the aircraft 5.

FIG. 6 shows a second embodiment of the inlet assembly including a rearsupport 101 that is designed to act as a firewall in the engineassembly. In the illustrated embodiment of FIG. 6, the rear support 101has a composite material outer radial portion 105 made of a heatresistant material (e.g., graphite composite) and a metallic innerradial portion 111 made of a deformable metal (e.g., aluminum).Alternatively, the rear support 101 could have a metallic inner radialportion and a metallic outer radial portion; a composite material innerradial portion and a composite material outer radial portion; or acomposite material inner radial portion and a metallic outer radialportion, without departing from the scope of this invention.

The outer radial portion 105 has ribs 107 and stiffeners (not shown)between the ribs 107 as in the first embodiment and the inner radialportion 111 includes energy absorbers 113 similar to the energyabsorbers 67 of the first embodiment. The ribs 107, energy absorbers113, and stiffeners are configured in a similar manner as describedabove for the first embodiment to form angularly spaced zones ofdeformation in the rear support 101. The inner radial portion 111 canhave an outer heat resistant layer covering the deformable metal to makethe inner portion of the rear support 101 resistant to heat. In theillustrated embodiment the inner radial portion 111 and the outer radialportion 105 are joined at respective overlapping portions 117, 119. Theoverlapping portions 117, 119 are attached by threaded fasteners (notshown) or other suitable methods (e.g., rivets, welding, etc.). The rearsupport 101 shown in FIG. 6 is angled towards the aft direction,however, the rear support may be otherwise positioned (e.g., vertical,angled in forward direction, etc.) without departing from the scope ofthis invention.

It is understood that zones of deformation of this embodiment functionin a similar manner as the first embodiment in that the rear support 101deforms at the zones of deformation during a fan blade-out event. Aswith the first embodiment, the rear support 101 is designed to deformwithout fracturing or shearing at the fasteners during the fan blade-outevent.

In view of the above, it will be seen that several advantageous resultsare obtained by the inlet assembly of the present invention. Forexample, the inlet assembly 17 having the rear support 35 is lightweightyet capable of withstanding the high stresses and strains during a fanblade-out event. The rear support 35 is designed to deform in the zoneof deformation 71 which acts as a crumple zone during a fan blade-outevent so that the rear support remains assembled in one piece withoutfracturing or breaking off during flight. Further, the rear support 101of the inlet assembly 17 may be designed to act as a firewall with anouter radial portion 105 made of heat resistant material such asgraphite composite and an inner radial portion 111 including the zone ofdeformation that is made of a deformable metal such as aluminum.

It will be further understood by those skilled in the art that while theforegoing has been disclosed above with respect to preferred embodimentsor features, various additions, changes, and modifications can be madeto the foregoing invention without departing from the spirit and scopethereof.

1. A method of manufacturing an inlet assembly for an aircraft enginenacelle having an outer barrel and an inner barrel, said methodcomprising: forming at least one radial stiffener in a rear support;forming at least one zone of deformation in said rear support by formingan arcuately extending energy absorber in said rear support, said energyabsorber being located in said rear support such that said support isdeformable in said zone of deformation in response to an applied forceduring a fan blade-out event to prevent fracture of said rear support;and attaching said rear support to the outer barrel and the inner barrelto form a closed axial end of the inlet assembly.
 2. The method of claim1 wherein forming said zone of deformation in said rear supportcomprises forming a plurality of zones of deformation angularly spacedaround said rear support.
 3. The method of claim 2 wherein attachingsaid rear support comprises attaching a plurality of annular sections ofsaid support to said inner and said outer barrel, each of said pluralityof annular sections comprising at least one zone of deformation of saidplurality of zones of deformation.
 4. The method of claim 1 whereinforming said energy absorber comprises forming a curved rib generallyadjacent an inner radial edge of said radial stiffener.
 5. The method ofclaim 4 wherein forming said radial stiffener comprises forming at leasttwo curved radial ribs at spaced-apart locations in said support and agenerally flat portion of said support between said at least two radialribs.
 6. The method of claim 5 wherein said rear support comprises aninner radial portion made of metal and said energy absorber is formed inthe inner radial portion.
 7. The method of claim 6 wherein said rearsupport comprises an outer radial portion made of a composite materialand said radial stiffener is formed in the outer radial portion.
 8. Themethod of claim 7 further comprising attaching said inner radial portionto said out radial portion.